Vane system with continuous support ring

ABSTRACT

A vane system includes a plurality of vane assemblies and a continuous support ring. The vane assemblies are arranged circumferentially about an axis and each include a hollow airfoil fairing and a spar. The spar has a spar flange and a spar leg extending radially inwardly from the spar flange and through the hollow airfoil fairing. The continuous support ring has radially inner and outer sides and defines a circumferential row of through-holes between the radially inner and outer sides. The spar legs extend through the through-holes and the spar flanges are affixed at the radially outer side of the continuous support ring.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.17/313,099 filed on May 6, 2021.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section may include low and high pressure compressors, andthe turbine section may also include low and high pressure turbines.

Airfoils in the turbine section are typically formed of a superalloy andmay include thermal barrier coatings to extend temperature capabilityand lifetime. Ceramic matrix composite (“CMC”) materials are also beingconsidered for airfoils. Among other attractive properties, CMCs havehigh temperature resistance. Despite this attribute, however, there areunique challenges to implementing CMCs in airfoils.

SUMMARY

A vane system according to an example of the present disclosure includesa plurality of vane assemblies arranged circumferentially about an axis.Each of the vane assemblies has a hollow airfoil fairing and a spar. Thespar has a spar flange and a spar leg that extends radially inwardlyfrom the spar flange and through the hollow airfoil fairing. Acontinuous support ring has radially inner and outer sides and defines acircumferential row of through-holes between the radially inner andouter sides. The spar legs extend through the through-holes, and thespar flanges are affixed at the radially outer side of the continuoussupport ring.

A further embodiment of any of the foregoing embodiments includes aplurality of fasteners extending through the spar flanges to theradially inner side of the continuous support ring and affixing the sparflanges to the continuous support ring.

In a further embodiment of any of the foregoing embodiments, each of thespars includes an additional spar leg that also extends from the sparflange and through the hollow airfoil fairing.

In a further embodiment of any of the foregoing embodiments, thecontinuous support ring includes a plurality of additionalthrough-holes. Each of the vane assemblies further include a second sparhaving a second spar flange and a second spar leg extending radiallyinwardly from the second spar flange and through the hollow airfoilfairing and one of the additional through-holes. The second spar flangeis affixed at the radially outward side of the continuous support ring.

A further embodiment of any of the foregoing embodiments includes atleast one cover plate radially constraining the spar flanges to affixthe spar flanges to the continuous support ring.

In a further embodiment of any of the foregoing embodiments, thecontinuous support ring includes a plurality of upstanding ridges on theradially outer side adjacent to the spar flanges. The plurality ofupstanding ridges axially and tangentially constraining the sparflanges.

In a further embodiment of any of the foregoing embodiments, the atleast one cover plate is a continuous ring.

In a further embodiment of any of the foregoing embodiments, the atleast one cover plate comprises a plurality of cover plates in acircumferential row. Each of the cover plates radially constraining oneof the spar flanges.

In a further embodiment of any of the foregoing embodiments, the atleast one cover plate comprises a first cover plate radiallyconstraining an axially forward end of the spar flange and a secondcover plate radially constraining an axially aft end of the spar flange.

In a further embodiment of any of the foregoing embodiments, thecontinuous support ring includes at least one wall extending in a radialdirection from the radially outer side thereof at an axially forward oran axially aft edge of the continuous support ring, and the at least onecover plate is affixed to the at least one wall.

A further embodiment of any of the foregoing embodiments includesfasteners extending through the spar flanges to the radially inner sideof the continuous support ring to further affix the spar flanges to thecontinuous support ring.

In a further embodiment of any of the foregoing embodiments, the sparflanges include at least one radially protruding lip at an axiallyforward end or axially aft end, the at least one radially protruding lipconfigured to mate with the at least one cover plate.

A gas turbine engine according to an example of the present disclosureincludes a compressor section, a combustor in fluid communication withthe compressor section, and a turbine section in fluid communicationwith the combustor. The turbine section has a vane system having aplurality of vane assemblies arranged circumferentially about an axis.Each of the vane assemblies has a hollow airfoil fairing and a spar. Thespar has a spar flange and a spar leg that extends radially inwardlyfrom the spar flange and through the hollow airfoil fairing. Acontinuous support ring has radially inner and outer sides and defines acircumferential row of through-holes between the radially inner andouter sides. The spar legs extend through the through-holes, and thespar flanges are affixed at the radially outer side of the continuoussupport ring.

A further embodiment of any of the foregoing embodiments includes anengine case surrounding the turbine section and wherein the continuoussupport ring is mounted to the engine case.

A further embodiment of any of the foregoing embodiments includes atleast one cover plate radially constraining the spar flanges to affixthe spar flanges to the continuous support ring.

In a further embodiment of any of the foregoing embodiments, the atleast one cover plate is integral with the engine case.

In a further embodiment of any of the foregoing embodiments, the atleast one cover plate comprises a first cover plate radiallyconstraining an axially forward end of the spar flange and a secondcover plate radially constraining an axially aft end of the spar flange,both of said first and second cover plates integral with the enginecase.

In a further embodiment of any of the foregoing embodiments, thecontinuous support ring includes a plurality of upstanding ridges on theradially outer side adjacent to the spar flanges. The plurality ofupstanding ridges axially and tangentially constraining the sparflanges.

In a further embodiment of any of the foregoing embodiments, thecontinuous support ring defines an annular planum, and each of the sparlegs includes a through-passage fluidly connected with the annularplanum.

A method for fabricating a vane system according to an example of thepresent disclosure includes providing a plurality of vane assembliesarranged circumferentially about an axis. Each of the vane assemblieshave a hollow airfoil fairing and a spar. The spar has a spar flange anda spar leg extending radially inwardly from the spar flange and throughthe hollow airfoil fairing. The method further includes providing acontinuous support ring having radially inner and outer sides anddefining a circumferential row of through-holes between the radiallyinner and outer sides. The spar legs are inserted through thethrough-holes and the spar flanges are affixed at the radially outerside of the continuous support ring.

The present disclosure may include any one or more of the individualfeatures disclosed above and/or below alone or in any combinationthereof.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 illustrates a gas turbine engine.

FIG. 2 illustrates a vane system from the engine.

FIG. 3 illustrates a vane assembly from the vane system.

FIG. 4 illustrates another example vane assembly with a spar that hastwo spar legs.

FIG. 5 illustrates another example vane assembly with a two spars.

FIG. 6 illustrates another example vane system with cover plates.

FIG. 7 illustrates a vane assembly from the vane system of FIG. 6 .

FIG. 8 illustrates a radial inward view of the vane assembly of FIG. 7 .

FIG. 9 illustrates a cross-sectional view of the example vane assemblyof FIG. 8 taken along line 9-9.

FIG. 10 illustrates another example vane assembly with a cover plate andfasteners.

FIG. 11 illustrates another example vane system with a cover plate thatis a continuous ring.

FIG. 12 illustrates another example vane system with a cover plate thatis continuous with a support structure.

FIG. 13 illustrates another example vane system with two cover platesthat are continuous with a support structure.

FIG. 14 illustrates a method of fabricating a vane system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a housing15 such as a fan case or nacelle, and also drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in the exemplary gas turbine 20 between thehigh pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 57 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded through the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), andcan be less than or equal to about 18.0, or more narrowly can be lessthan or equal to 16.0. The geared architecture 48 is an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3. The gear reduction ratio maybe less than or equal to 4.0. The low pressure turbine 46 has a pressureratio that is greater than about five. The low pressure turbine pressureratio can be less than or equal to 13.0, or more narrowly less than orequal to 12.0. In one disclosed embodiment, the engine 20 bypass ratiois greater than about ten (10:1), the fan diameter is significantlylarger than that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about five 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to aninlet of low pressure turbine 46 as related to the pressure at theoutlet of the low pressure turbine 46 prior to an exhaust nozzle. Thegeared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1 and less than about 5:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (′ TSFC′)” is the industry standard parameter of 1 bm offuel being burned divided by 1 bf of thrust the engine produces at thatminimum point. The engine parameters described above and those in thisparagraph are measured at this condition unless otherwise specified.“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45, or more narrowly greater than orequal to 1.25. “Low corrected fan tip speed” is the actual fan tip speedin ft/sec divided by an industry standard temperature correction of[(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” asdisclosed herein according to one non-limiting embodiment is less thanabout 1150.0 ft/second (350.5 meters/second), and can be greater than orequal to 1000.0 ft/second (304.8 meters/second).

FIG. 2 illustrates a vane system 60 from the turbine section 28 of theengine 20. Although the vane system 60 is shown and described withreference to application in the turbine section 28, it is to beunderstood that the examples herein are also applicable to structuralvanes in other sections of the engine 20.

The vane system 60 is comprised of a continuous support ring 62 and aplurality of vane assemblies 64 arranged circumferentially about theengine central longitudinal axis A. Each vane assembly 64 includes ahollow airfoil fairing 66 and a spar 68 that is attached to thecontinuous support ring 62 to support the hollow airfoil fairing 66.

FIG. 3 illustrates an example of one of the vane assemblies 64. Thehollow airfoil fairing 66 includes several sections, including anairfoil section 70, and first and second platforms 72/74, between whichthe airfoil section 70 extends. The airfoil section 70 circumscribes aninterior through-cavity 76 and generally extends in a radial directionrelative to the central engine longitudinal axis A. The terms such as“inner” and “outer” refer to location with respect to the central engineaxis A, i.e., radially inner or radially outer. Moreover, theterminology “first” and “second” used herein is to differentiate thatthere are two architecturally distinct components or features. It is tobe further understood that the terms “first” and “second” areinterchangeable in that a first component or feature could alternativelybe termed as the second component or feature, and vice versa.

The hollow airfoil fairing 66 is continuous in that the platforms 72/74and the airfoil section 70 constitute a unitary body. As an example, theairfoil fairing 66 is formed of a ceramic matrix composite, an organicmatrix composite (OMC), or a metal matrix composite (MMC). For instance,the ceramic matrix composite (CMC) is formed of ceramic fiber tows thatare disposed in a ceramic matrix. The ceramic matrix composite may be,but is not limited to, a SiC/SiC ceramic matrix composite in which SiCfiber tows are disposed within a SiC matrix. Example organic matrixcomposites include, but are not limited to, glass fiber tows, carbonfiber tows, and/or aramid fiber tows disposed in a polymer matrix, suchas epoxy. Example metal matrix composites include, but are not limitedto, boron carbide fiber tows and/or alumina fiber tows disposed in ametal matrix, such as aluminum. A fiber tow is a bundle of filaments. Asan example, a single tow may have several thousand filaments. The towsmay be arranged in a fiber architecture, which refers to an orderedarrangement of the tows relative to one another, such as, but notlimited to, a 2D woven ply or a 3D structure.

The spar 68 mechanically supports the hollow airfoil fairing 66. Thespar 68 in this example includes a spar flange 68 a and a spar leg 68 bthat extends from the spar flange 68 a. The spar leg 68 b extendsthrough the through-cavity 76 and protrudes from the platform 74 of theairfoil fairing 66. The end of the spar leg 68 b is received throughsupport hardware 78 and is secured thereto by an attachment, such as apin, so as to trap the airfoil fairing 66 between the support hardware78 and the spar flange 68 a.

The spar leg 68 b defines one or more interior through-passages 68 c.Cooling air, such as bleed air from the compressor section 24, isconveyed into and through the through-passage 68 c of the spar 68. Thiscooling air may be destined for a downstream cooling location, such as atangential onboard injector (TOBI). Cooling air may also be providedinto the cavity 76 for cooling of the airfoil section 70.

The continuous support ring 62 is secured to a fixed support structure(not shown), such as engine static structure 36 or an engine case. Thecontinuous support ring 62 is axisymmetric and extends endlessly aroundthe central engine axis A. The continuous support ring 62 has a radiallyinner and outer sides 62 a/62 b and a circumferential row ofthrough-holes 80 that extend between the radially inner and outer sides62 a/62 b. The continuous support ring 62 may be formed of a metalalloy, such as a nickel- or cobalt-based alloy. In one example, thecontinuous support ring 62 is formed of a single, monolithic body thatis free of any mechanical joints or seams, such as weld joints. In otherexamples, however, the continuous support ring 62 may be formed of arcsegments that are welded or otherwise metallurgically attached togetherto form one continuous ring. Thus, the continuous support ring 62 isfree of any inter-segment spaced joints.

The continuous support ring 62 further includes walls 79 that extend ina radial direction from the radially outer side 62 b at an axiallyforward edge 62 c and an axially aft edge 62 d. The walls 79 andradially outer side 62 b of the continuous support ring 62 define threesides of an annular plenum 63, which may be used to deliver cooling air,illustrated as arrow D, to the through-passages 68 c of the spar 68. Theannular plenum 63 may further be enclosed by the fixed support structurethat the continuous support ring 62 is attached to.

As illustrated in FIG. 3 , the spar leg 68 b extends radially inwardlythrough the through-hole 80. The spar flange 68 a is affixed at theradially outer side 62 b of the continuous support ring 62 by aplurality of fasteners 82 that extend through the spar flanges 68 a andthe continuous support ring 62 to the radially inner side 62 a of thecontinuous support ring 62. The fasteners 82 may be, but are not limitedto, screws, rivets, or bolts. Alternatively, or in combination with thefasteners 82, the spar flange 68 a may be welded to the outer side 62 bof the continuous support ring 62 at a weld joint 83.

The continuous support ring 62 facilitates robust support of the vaneassemblies 64 because it is free of inter-segment spaced joints, andthus, is relatively stiff. Further, the absence of inter-segment spacedjoints facilitates assembly and manufacturing of the vane system 60, andalso facilitates easier positioning of the vane system 60 and affixtureof vane system 60 to engine case 36 because there are fewer pieces tomanufacture and install. The configuration also facilitates sealing ofthe annular plenum 63 for delivery of the cooling air to thethrough-passages 68 c because there are fewer potential leak pathswithout inter-segment spaced joints.

FIG. 4 illustrates another example vane assembly 164 for the vane system60. In this disclosure, like reference numerals designate like elementswhere appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding elements.

In the vane assembly 164 the spar 168 includes two spar legs 168 b/168 dthat extend from the spar flange 168 a and through the hollow airfoilfairing 66. The spar legs 168 b/168 d may both extend through a singlethrough-hole 80 of the continuous support ring 62, or the continuoussupport ring 62 may include two through-holes 80 such that each spar leg168 b/168 d extends through a respective through hole 80. The use ofmultiple spar legs 168 b/168 d may facilitate a more robust mechanicalsupport of the airfoil fairing 66. Although two spar legs 168 b/168 dare shown, the spar 168 may include additional spar legs for addedstiffness and support of the airfoil fairing 66 if space allows.

FIG. 5 illustrates another example vane assembly 264. The vane assembly264 in this example includes a first spar 268 and a second spar 269,each with a respective spar flange 268 a/269 a and spar leg 268 b/269 b.The spar legs 268 b/269 b extend radially inwardly from the spar flanges268 a/269 a, through the through-holes 80 of the continuous support ring62, and through the hollow airfoil fairing 66. The spar flanges 268a/269 a of the first and second spars 268/269 are both affixed at theradially outer side 62 b of the continuous support ring 62 by fasteners82. Although two spars 268/269 are shown, additional spars may beincluded and affixed to the continuous support ring 62 if desired andspace allows.

The use of the two spars 268/269 may facilitate assembly of the vaneassembly 264. For instance, if an airfoil section has an extreme bowand/or twist or there is a tight fit with two spar legs that areattached to a common spar flange, insertion of the spar legs into theairfoil section may not be possible due to interference. However, byhaving two separate spars 268/269, each one can be independentlymanipulated for insertion without constraint by the other.

FIG. 6 illustrates another example vane system 160 in which a pluralityof cover plates 84 affix the vane assemblies 64 to the continuoussupport ring 62 instead of fasteners as in the previous examples. Asshown in FIG. 7 , each of the cover plates 84 includes a radial leg 84 aand an axial leg 84 b that form an L-shape. The radial leg 84 a isaffixed to the wall 79 of the continuous support ring 62 by thefasteners 82 that extend through the radial leg 84 a and the wall 79.The axial leg 84 b clamps the spar flange 68 a to the radially outerside 62 b of the continuous support ring 62 so as to radially constrainthe spar flange 68 a. In this example, there are two cover plates 84 at,respectively, an axially forward end 68 e of the spar flange 68 a and anaxially aft end 68 f of the spar flange 68 a. Thus, each set of twocover plates 84 radially constrains one spar flange 68 a. Use ofdiscrete sets of two cover plates 84 for each flange 68 a allows theclamping force and positioning of each cover plate 84 to be individuallyadjusted for each spar flange 68 a, which may facilitate tailoring ofthe configuration to account for dimensional tolerances.

FIG. 8 illustrates a radial inward view of the vane system 160,excluding the cover plates 84. The continuous support ring 62 includes aplurality of upstanding ridges 88 on the radially outer side 62 b. Theupstanding ridges 88 are at positions adjacent to the spar flange 68 a,such that the upstanding ridges 88 axially and tangentially constrainmovement of the spar flange 68 a. In this example, a first upstandingridge 88 a and second upstanding ridge 88 b are elongated in an axialdirection and abut a first tangential end 68 g and second tangential end68 h of the spar flange 68 a respectively. A third upstanding ridge 88 cincludes an elbow such that it abuts both first tangential end 68 g andaxially forward end 68 e of spar flange 68 a. It is to be appreciatedthat the aerodynamic loads on the airfoil fairing 66 that aretransmitted to the spar 68 may tend to always move the spar 68 (and thusthe spar flange 68 a) in the same direction. In this regard, thelocations of the upstanding ridges 88 may be selected in accordance withthe directional movement tendency of the spar platform 68 a for thegiven aerodynamic loads.

FIG. 9 illustrates a cross-sectional view of vane system 160 taken alongline 9-9 of FIG. 8 (but with the cover plate 84). The cover plates 84are rigid components with a radially inner face 84 c facing the sparflange 68 a. The radially inner face 84 c is positioned radiallyoutwards of the spar flange 68 a and mates with the spar flange 68 a asan interference fit. Accordingly, the spar flange 68 a abuts the coverplates 84 and is thereby constrained from moving radially, while theupstanding ridges 88 constrain tangential and axial movement.

FIG. 10 illustrates an example vane system 260 that is a combination ofthe examples of FIGS. 3 and 7 . Here, the cover plate 84 radiallyconstrains the axially forward end 68 e of the spar flange 68 a, and thefasteners 82 extend through the axially aft end 68 f of the spar flange68 a to the radially inner side 62 a of the continuous support ring 62.Alternatively, the cover plate 84 may radially constrain the axially aftend 68 f of the spar flange 68 a and the fasteners 82 may extend throughthe axially forward end 68 e.

FIG. 11 illustrates an example vane system 360 that is similar to theexample in FIG. 6 except that the cover plates 184 are continuous ringsthat radially constrain all of the spar flanges 68 a of the plurality ofvane assemblies 64 on continuous support ring 62. In this example, onecover plate 184 radially constrains the axially forward end 68 e of thespar flanges 68 a and another cover plate 184 radially constrains theaxially aft end 68 f of the spar flanges 68 a. The use of continuousring cover plates 184, rather than the plurality of discrete coverplates 84 as in FIG. 10 , eliminates the use of multiple pieces andthereby facilitates assembly and manufacturing of vane system 360.However, while assembly may be easier, the continuous ring cover plates184 cannot be tailored to the individual tolerances of the spar flanges68 a as the discrete cover plates 84 can.

FIG. 12 illustrates an example vane system 460 including a first coverplate 90 and second cover plate 84 both radially constraining the sparflange 368 a of the vane assembly 364 on the continuous support ring 62.In this example, the first cover plate 90 is integral with a supportstructure 92, which may be attached to, or integral with, the enginecase 36. For example, the first cover plate 90 is configured as anextension arm that protrudes from the support structure 92. Fasteners 82attach the first cover plate 90 to the continuous support ring 62, andalso attach vane system 460 to the support structure 92. Such aconfiguration reduces the number of components by eliminating some ofthe prior described cover plates 84.

As shown in FIG. 12 , the spar flange 368 a also includesradially-protruding lips 94 at both axially forward and aft ends 368e/368 f of the spar flange 368 a. The geometry of the lips 94 match thegeometry of the under-surfaces the cover plates 84/90 such that the lips94 make area contact with the cover plates 84/90 (as opposed to point orline contact). The area contact facilitates load distribution to therebyenhance constraint of the spar flanges 368 a.

FIG. 13 illustrates an example vane system 560 that is similar to theexample in FIG. 12 except that both a first cover plate 190 and a secondcover plate 191 are integral with a support structure 192, such as anengine case. The first and second cover plates 190/191 extend radiallyinwardly from the support structure 192 to radially constrain the sparflange 468 a of the vane assembly 464 on the continuous support ring 62.The axially forward end 468 e of the spar flange 468 a includes aradially outwardly extending wall 96 that interfaces with the secondcover plate 191. In the illustrated example, the support structure 192includes case sections 192 a/192 b that each have a radial flange 192 c.In this example, a wall 79 of the continuous support ring 62 issandwiched axially between the flanges 192 c and a fastener 82 extendsthrough the flanges 192 c and the wall 79 to secure the case sections192 a/192 b together and also attach vane system 560 to the supportstructure 192. Similar to the example in FIG. 12 , such a configurationfurther reduces the number of components in the system.

Referring back to FIGS. 2-5 , the example vane system 60 usesradially-oriented fasteners 82 to secure the spar flanges 68 a to thecontinuous support ring 62. However, there may be spatial limitationswhen installing a vane system that limits the use of suchradially-oriented fasteners. For example, there may not be space betweenthe platform 72 and the continuous support ring 62 for theradially-oriented fasteners 82 to extend through the radially inner side62 a of the continuous support ring 62. If such spatial limitations arepresent, the example vane systems 160, 260, 360, 460, and 560illustrated in FIGS. 6-13 may be used to eliminate the fasteners andinstead use the cover plates, which do not require space between theplatform 72 and the continuous support ring 62.

FIG. 14 illustrates a method 500 of fabricating the vane system 60described above. At steps 502 and 504 the vane assemblies 64 and thecontinuous support ring 62 are provided. Most typically, the vaneassemblies 64 and the continuous support ring 62 are furnished aspre-fabricated components. However, the provision of the vane assemblies64 and/or the continuous support ring 62 may alternatively involvefabrication of one or more of the components of the vane assemblies 64and/or the continuous support ring 62.

Step 506 includes inserting the spar legs 68 b through the through-holes80 of the continuous support ring 62. The insertion may involve, eithermanually or through automation, aligning the spar legs 68 b withthrough-holes 80 and moving the spar legs 68 b through the through-holes80 until the spar flanges 68 a contact the radially outer side 62 b ofthe continuous support ring 62. Step 508 includes affixing the sparflanges 68 a at the radially outer side 62 b of the continuous supportring 62. For the example vane system 60, the affixing involves, eithermanually or through automation, inserting the fasteners 84 through thespar flange 68 a and the continuous support ring 62 and tightening thefasteners 84. For the example vane systems 160, 260, 360, and 460, theaffixing involves providing a cover plate 84, and then, either manuallyor through automation, positioning the cover plate 84 radially over theflanges 68 a, inserting fasteners 82 through the cover plate 84 and awall 79 of the continuous support ring 62, and tightening the fasteners82. Although method 500 has been described with reference to the vanesystem 60, it is to be understood that the method 500 is also applicableto the vane systems 160, 260, 360, 460, and 560 described above. For theexample vane system 560, the affixing involves, either manually orthrough automation, providing a case section 192 a that has integralcover plates 190/191, positioning the support structure 192 such thatthe cover plates 190/191 are radially over the flanges 68 a, inserting afastener 82 through the case section 192 a, a wall 79 of the continuoussupport ring 62, and a second case section 192 b of the supportstructure 192, and tightening the fastener 82. Further, the vane systemsof this disclosure may be dissembled via the reverse procedure of method500 for repair, component replacement, maintenance, or the like.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthis disclosure. The scope of legal protection given to this disclosurecan only be determined by studying the following claims.

What is claimed is:
 1. A vane system comprising: a plurality of vaneassemblies arranged circumferentially about an axis, each of the vaneassemblies including: a hollow airfoil fairing, and a spar, the sparhaving a spar flange and a spar leg extending radially inwardly from thespar flange and through the hollow airfoil fairing; and a continuoussupport ring having radially inner and outer sides and defining acircumferential row of through-holes between the radially inner andouter sides, the spar legs extending through the through-holes and thespar flanges being affixed at the radially outer side of the continuoussupport ring.
 2. The vane system of claim 1, further comprising aplurality of fasteners extending through the spar flanges to theradially inner side of the continuous support ring and affixing the sparflanges to the continuous support ring.
 3. The vane system of claim 1,wherein each of the spars includes an additional spar leg that alsoextends from the spar flange and through the hollow airfoil fairing. 4.The vane system of claim 1, wherein the continuous support ring includesa plurality of additional through-holes, and wherein each of the vaneassemblies further comprises a second spar having a second spar flangeand a second spar leg extending radially inwardly from the second sparflange and through the hollow airfoil fairing and one of the additionalthrough-holes, the second spar flange being affixed at the radiallyoutward side of the continuous support ring.
 5. The vane system of claim1, further comprising at least one cover plate radially constraining thespar flanges to affix the spar flanges to the continuous support ring.6. The vane system of claim 5, wherein the continuous support ringincludes a plurality of upstanding ridges on the radially outer sideadjacent to the spar flanges, the plurality of upstanding ridges axiallyand tangentially constraining the spar flanges.
 7. The vane system ofclaim 5, wherein the at least one cover plate is a continuous ring. 8.The vane system of claim 5, wherein the at least one cover platecomprises a plurality of cover plates in a circumferential row, each ofthe cover plates radially constraining one of the spar flanges.
 9. Thevane system of claim 5, wherein the at least one cover plate comprises afirst cover plate radially constraining an axially forward end of thespar flange and a second cover plate radially constraining an axiallyaft end of the spar flange.
 10. The vane system of claim 5, wherein thecontinuous support ring includes at least one wall extending in a radialdirection from the radially outer side thereof at an axially forward oran axially aft edge of the continuous support ring, and the at least onecover plate is affixed to the at least one wall.
 11. The vane system ofclaim 5, further comprising fasteners extending through the spar flangesto the radially inner side of the continuous support ring to furtheraffix the spar flanges to the continuous support ring.
 12. The vanesystem of claim 5, wherein the spar flanges include at least oneradially protruding lip at an axially forward end or axially aft end,the at least one radially protruding lip configured to mate with the atleast one cover plate.
 13. A gas turbine engine comprising: a compressorsection; a combustor in fluid communication with the compressor section;and a turbine section in fluid communication with the combustor, theturbine section having a vane system including: a plurality of vaneassemblies arranged circumferentially about an axis, each of the vaneassemblies including: a hollow airfoil fairing, and a spar, the sparhaving a spar flange and a spar leg extending radially inwardly from thespar flange and through the hollow airfoil fairing; and a continuoussupport ring having radially inner and outer sides and defining acircumferential row of through-holes between the radially inner andouter sides, the spar legs extending through the through-holes and thespar flanges being affixed at the radially outer side of the continuoussupport ring.
 14. The gas turbine engine of claim 13, further comprisingan engine case surrounding the turbine section and wherein thecontinuous support ring is mounted to the engine case.
 15. The gasturbine engine of claim 14, further comprising at least one cover plateradially constraining the spar flanges to affix the spar flanges to thecontinuous support ring.
 16. The gas turbine engine of claim 15, whereinthe at least one cover plate is integral with the engine case.
 17. Thegas turbine engine of claim 16, wherein the at least one cover platecomprises a first cover plate radially constraining an axially forwardend of the spar flange and a second cover plate radially constraining anaxially aft end of the spar flange, both of said first and second coverplates integral with the engine case.
 18. The gas turbine engine ofclaim 15, wherein the continuous support ring includes a plurality ofupstanding ridges on the radially outer side adjacent to the sparflanges, the plurality of upstanding ridges axially and tangentiallyconstraining the spar flanges.
 19. The gas turbine engine of claim 14,wherein the continuous support ring defines an annular planum, and eachof the spar legs includes a through-passage fluidly connected with theannular planum.
 20. A method for fabricating a vane system, the methodcomprising: providing a plurality of vane assemblies arrangedcircumferentially about an axis, each of the vane assemblies including:a hollow airfoil fairing, and a spar, the spar having a spar flange anda spar leg extending radially inwardly from the spar flange and throughthe hollow airfoil fairing; providing a continuous support ring havingradially inner and outer sides and defining a circumferential row ofthrough-holes between the radially inner and outer sides; inserting thespar legs through the through-holes; and affixing the spar flanges atthe radially outer side of the continuous support ring.